Q1) Air at mach number 2 over a wedge and generation of an oblique shock wave 40° . determine : 1. Static pressure and temperature downstream the shock. 2. The mach number of the flow downstream the shock. , pressure 0.5* 10 N/m² and a temperature of 0 °C , 3. The wedge angle. ********* ********* ******** ***********************

Solve plzzz asap... Habd written

Q2/ Answer the following: 1) Separation in flow past a solid object is caused by (a) a favourable (negative ) pressure gradient (b) an adverse ( positive ) pressure gradient (c) the boundary layer thickness reducing to zero 2) In the design, we have to delay the separation point in order to reduce wake size (T/F) 3) Define Mach number, boundary layer and shock wave. 4) when the flow is accelerates if the cross sectional area is increases (a) M 1 (c) M =1

Consider steady, isentropic flow of air through a duct with varying area. At section 1, p1 = 15 kPa (abs), T = 10°C, Vị = 592 m/s, A1 = 0.25 m². At a downstream section 2, T2 = 137°C, Maz = 0.75. Determine: a. Mach number, Mai b. sketch geometry of the duct c. mass flow rate [kg/s] d. pressure, P2 [kPa] e. cross-sectional area, A2 [m2]

!

Q2: The large compressed-air tank shown in the figure bellow exhausts from a nozzle at an exit velocity of V235 m/s. Assuming isentropic flow, compute: a) the exit Mach number. b) the pressure in the tank. (take; k = 1.396, C, = 917 J/Kg.k and R = 260 J/Kg.k) air at 30 C tank conditions Pan= 101 LPa

Q2: Air (y = 1.4) enters a converging-diverging diffuser with a Mach number of 2.8, static pressure pi of 100 kPa, and a static temperature of 20°C. For the flow situation shown in Figure, find the exit velocity, exit static pressure, and exit stagnation pressure. Ans: Ve = 55.093 m/s; Pe = 2236.678 kPa; Poe = 2252.44 kPa i M₁ = 2.8 A₁ = 0.10 m² A₂ = 0.50 m² A₁ = 0.25 m²

1. The flow is considered compressible when the value of Mach number is.... (A) greater than unity (B) greater or equal to 0.3 (C) greater than 0.5 (D) less than 0.3 2. In c-d nozzle, when dA/dV is positive, then the flow is.... (A) subsonic (B) sonic (C) supersonic (D) incompressible 3. Along the adiabatic duct, the...... remains constant. (A) stagnation pressure (B) static pressure (C) static temperature (D) stagnation temperature 4. Across the normal shock wave, the gas velocity ...... (A) decreases↓ (B) increases ↑ (C) remains constant (D) varies randomly v2 (C) 2Cp (D) PV² 5. The dynamic temperature can be calculated from..... (A): M² 2Cp (B) -M² B 6. A supersonic flow with 38° oblique shock angle passes over a concave corner of angle 0=20°. Then, the free stream Mach number M₁ is.... (A) 2 (B) 5 7. Across the expansion waves, the ..... remains constant. (A) Mach number (B) stagnation pressure (C) 3 (C) gas velocity (D) 4 (D) temperature. (D) Pe=0 8. The maximal thrust of the…

answer quickly

1.AB